The present invention relates generally to gas turbine engines, and, more specifically, to cooling of turbine rotor blades and stator vanes therein.
A gas turbine engine includes a compressor that compresses air which is channeled to a combustor wherein it is mixed with fuel and ignited for generating combustion gases. The combustion gases flow downstream through one or more stages of turbines which extract energy therefrom for powering the compressor and producing additional output power for driving a fan for powering an aircraft in flight for example.
A turbine stage includes a row of turbine rotor blades secured to the outer perimeter of a rotor disk, with a stationary turbine nozzle having a plurality of stator vanes disposed upstream therefrom. The combustion gases flow between the stator vanes and between the turbine blades for extracting energy to rotate the rotor disk.
Since the combustion gases are hot, the turbine vanes and blades are typically cooled with a portion of compressor air bled from the compressor for this purpose. Diverting any portion of the compressor air from use in the combustor necessarily decreases the overall efficiency of the engine.
Accordingly, it is desired to cool the vanes and blades with as little compressor bleed air as possible.
Turbine vanes and blades include an airfoil over which the combustion gases flow. The airfoil typically includes one or more serpentine cooling passages therein through which the compressor bleed air is channeled for cooling the airfoil. The airfoil may include various turbulators therein for enhancing cooling effectiveness, and the cooling air is discharged from the passages through various film cooling holes disposed around the outer surface of the airfoil.
The airfoil outer surface is defined by a generally concave pressure side and an opposite, generally convex suction side which extend radially between a root and a tip of the airfoil and axially between leading and trailing edges thereof. The temperature profiles of the combustion gases channeled over the airfoil vary significantly over the pressure and suction sides. This in turn affects both the cooling requirements over the airfoil and cooling effectiveness. Greater cooling is desired where heat input is greatest, and backflow margin and blowing ratio must be controlled across the film cooling holes. Film cooling holes should have suitable blowing ratios to most effectively produce a protecting layer of film cooling air over the blade surface without flow separation and with suitable backflow margin.
In U.S. Pat. No. 5,591,007-Lee et al, a multi-tier turbine airfoil is disclosed and claimed for improving cooling performance of internal serpentine cooling passages therein. By arranging two or more serpentine passages in independent circuits radially over the airfoil span, cooling air may be directly channeled to the mid-span portion of the airfoil having the greatest heat input from the combustion gases. Since the cooling air in each serpentine circuit increases in temperature as it cools the airfoil, the multi-tier serpentine circuits target the cooling air to specific regions of the airfoil for enhancing cooling thereof over conventional radial serpentine circuits extending completely between the root and tip of the airfoil.
However, in order to effect the multi-tier serpentine circuit, radially extending span ribs must necessarily be interrupted for providing a separating tier rib extending axially. Since a turbine rotor blade experiences substantial centrifugal force during operation, the interruption in the span rib interrupts the loadpath for carrying airfoil centrifugal loads to the dovetail and in turn to the rotor disk supporting the airfoil.
The centrifugal loads must then be carried solely through respective portions of the airfoil pressure and suction sides bypassing the span interruption which results in local stress concentration that affects the useful life of the airfoil. Furthermore, an axially extending tier partition rib effectively introduces mere dead weight without significant load support, and the centrifugal loads therefrom must also be carried by the adjacent portions of the airfoil.
Accordingly, although the multi-tier turbine airfoil provides improvement in cooling thereof, it also changes the load carrying structure of the airfoil for both centrifugal forces, as well as vibratory response. It is therefore desired to further improve the multi-tier turbine airfoil both in load carrying capability and cooling effectiveness.